Gas turbine engines typically include a compressor, a combustor, and a turbine, with an annular flow path extending axially through each. Initially, air flows through the compressor where it is compressed or pressurized. The combustor then mixes and ignites the compressed air with fuel, generating hot combustion gases. These hot combustion gases are then directed from the combustor to the turbine where power is extracted from the hot gases by causing blades of the turbine to rotate.
A thermal management system in the gas turbine engine may maintain operable temperatures for fuel, oil, and other fluids communicated throughout the engine. Typically, thermal management systems may include one or more heat exchangers for transferring heat between the various fluids. The thermal management system may manage heat generated by the engine during operation. For example, the thermal management system may communicate conditioned fluids to various systems in order to minimize heat generation and dissipate the heat. Heat may be transferred into the engine fuel in order to increase fuel efficiency and engine performance.
However, typical thermal management systems have not attempted to protect hardware in the fuel system. Fuel system hardware in gas turbine engines may experience contamination and lacquer build-up. For example, in aircraft applications, fuel lacquering or other deposit formation may be induced at low altitudes due to the high oxygen content of fuel and elevated fuel temperatures. Accordingly, there exists a need for a thermal management system and method that provides fuel efficiency, while protecting against deposit formation on the fuel hardware. This invention is directed to solving this need and others.